This invention relates generally to turbomachinery compressors and more particularly relates to rotor blade stages of such compressors.
A gas turbine engine includes, in serial flow communication, a compressor, a combustor, and turbine. The turbine is mechanically coupled to the compressor and the three components define a turbomachinery core. The core is operable in a known manner to generate a flow of hot, pressurized combustion gases to operate the engine as well as perform useful work such as providing propulsive thrust or mechanical work. One common type of compressor is an axial-flow compressor with multiple rotor stages each including a disk with a row of axial-flow airfoils, referred to as compressor blades.
For reasons of thermodynamic cycle efficiency, it is generally desirable to incorporate a compressor having the highest possible pressure ratio (that is, the ratio of inlet pressure to outlet pressure). It is also desirable to include the fewest number of compressor stages. However, there are well-known inter-related aerodynamic limits to the maximum pressure ratio and mass flow possible through a given compressor stage.
It is known to reduce weight, improve rotor performance, and simplify manufacturing by minimizing the total number of compressor airfoils used in a given rotor blade row. However, as airfoil blade count is reduced the accompanying reduced hub solidity tends to cause the airflow in the hub region of the rotor airfoil to undesirably separate from the airfoil surface.
It is also known to configure the disk with a non-axisymmetric “scalloped” surface profile to reduce mechanical stresses in the disk. An aerodynamically adverse side effect of this feature is to increase the rotor blade row through flow area and aerodynamic loading level promoting airflow separation.
Accordingly, there remains a need for a compressor rotor that is operable with sufficient stall range and an acceptable balance of aerodynamic and structural performance.